KR20140049548A - Alloy, protective layer and component - Google Patents

Abstract

Known protective layers containing high Cr content and further silicon form embrittlement phases which are further embrittled under the influence of carbon during use. The protective layer according to the invention comprises 22% to 24% cobalt (Co), 10.5% to 11.5% aluminum (Al), 0.2% to 0.4% yttrium (Y), and / or elements of scandium and rare earths. One or more equivalent metals selected from the group comprising: 14% to 16% of chromium (Cr), optionally 0.3% to 0.9% of tantalum, nickel (Ni).

Description

{ALLOY, PROTECTIVE LAYER AND COMPONENT}

The invention relates to an alloy according to claim 1, to a protective layer for protecting the part against corrosion and / or oxidation, particularly at high temperatures, according to claim 13, and to a part according to claim 15.

Protective layers for metal parts for which corrosion and / or oxidation resistance should be improved are well known in the art. Most of the protective layers are known under the aggregate noun MCrAlY, where M represents one or more of the elements selected from the group comprising iron, cobalt and nickel and further major components are chromium, aluminum and yttrium.

Typical coating layers of this type are known from US Pat. Nos. 4,005,989 and 4,034,142.

In the case of stationary gas turbines, as well as for aircraft engines, the effort to raise the inlet temperature is significant in the gas turbine specialization because it is an important parameter for the thermodynamic efficiency attained by the gas turbine. Because it is. With the use of specially developed alloys as the base material for parts to be exposed to high thermal loads, such as guide vanes and rotor blades, inlet temperatures well above 1000 ° C. are possible, in particular through the use of single crystal superalloys. In the meantime, the prior art allows not only temperatures above 950 ° C. for stationary gas turbines, but also inlet temperatures above 1100 ° C. for gas turbines of aircraft engines. Examples of the construction of a turbine blade comprising a single crystal substrate, which can itself be complicated, are disclosed in WO 91/01433 A1.

The physical load carrying capacity of the base materials developed so far for components exposed to high loads is practically no problem with possible further rises in inlet temperature, while the protective layers can be used to achieve sufficient resistance to oxidation and corrosion. Must be used. In addition to the sufficient chemical resistance of the protective layer under the action expected by the flue gas at temperature conditions on the order of 1000 ° C., the protective layer must also possess sufficiently good mechanical properties, especially with regard to the mechanical interaction between the protective layer and the base material. In particular, since the working points for oxidation and corrosion can be provided in this way, the protective layer must have sufficient ductility, in order to cause deformation of the base material and thus not to rupture. In this case, typically, an increase in the proportion of elements such as aluminum and chromium, which can improve the resistance of the protective layer to oxidation and corrosion, deteriorates the ductility of the protective layer and, therefore, at the mechanical load conditions normally encountered in gas turbines. The problem arises that mechanical failures, in particular the formation of cracks, must be taken into account.

It is therefore an object of the present invention to provide an alloy and a protective layer which have excellent high temperature resistance during corrosion and oxidation and have excellent long-term stability but are particularly suitably matched to the mechanical loads expected at high temperatures, especially in gas turbines.

This problem is solved by the alloy according to claim 1 and the protective layer according to claim 10.

A further object of the present invention is to provide a component comprising an improved protection against corrosion and oxidation. This problem is likewise solved by a part according to claim 12, in particular by a part of a gas turbine or a steam turbine, comprising a protective layer of the type described above for protection against corrosion and oxidation at high temperatures.

The dependent claims enumerate further preferred measures which may optionally be combined with one another in the desired type and manner.

The invention is explained in more detail below.

The drawings and specification only show embodiments of the invention.

1 shows a layer system comprising a protective layer.
2 is a diagram describing the composition of superalloys.
3 shows a gas turbine.
4 shows a turbine blade.
5 shows a combustion chamber.

According to the invention, the protective layer 7 (FIG. 1) for protecting the part against corrosion and oxidation at high temperatures consists essentially of the following elements (ratio units: weight%).

nickel,

Co: 22% ~ 24%,

Cr: 14%-16%,

Al: 10.5%-11.5%,

0.2% to 0.4% rare earth elements (yttrium, ...) and / or scandium (Sc):

Optionally

Ta: 0.3% to 0.9%.

The enumerated list of alloying elements Ni, Co, Cr, Al, Y, Ta is preferably not definite.

Nickel preferably forms a matrix.

Preferably, the enumerated list of Ni, Co, Cr, Al, Y, Ta is definite.

The content of alloying elements Co, Cr, Al, Y has the following advantages.

Co content above medium:

Enlargement of the beta / gamma field, eg prevention of embrittlement such as alpha phase.

Medium Cr Content:

High enough to increase the activity of Al for Al ₂ O ₃ formation;

Low enough to prevent embrittlement phase (alpha-chrome or sigma phase).

Al content above medium:

High enough for Al activity to form a stable Al ₂ O ₃ layer;

Low enough to prevent embrittlement effects.

Low Y content:

High enough to form Y-aluminate for the formation of Y-containing "pegs" at low oxygen contamination;

Low enough to negatively promote oxide layer growth of the Al ₂ O ₃ layer.

Tantalum has a positive effect on the phase stability of the γ 'phase or delays the transition to higher temperatures, thus slowing phase degradation through the consumption of aluminum in the layer.

It is noted here that the proportion of the individual elements is particularly adjusted in relation to its action which is found in particular with respect to the element of silicon. If the proportion is assigned such that no silicon precipitate is formed, preferably no embrittlement occurs during use of the protective layer, whereby the operating time behavior is improved and extended. This is achieved not only by the low chromium content, but also by the correct assignment of the aluminum content, taking into account the influence of aluminum on the phase formation. Especially in the interaction with the reduction of the embrittlement phase which acts negatively under higher mechanical properties, the mechanical properties are improved through the reduction of the mechanical stress by the selected nickel content.

The protective layer is excellent in corrosion resistance but has particularly good resistance to oxidation and is characterized by particularly good ductility properties, whereby the protective layer is applied in the gas turbine 100 (FIG. 3) under conditions of further elevated inlet temperature. Particularly suitable for Embrittlement rarely occurs during operation because the layer contains very little chromium-silicon precipitate that is embrittled during use.

The powder is applied by, for example, plasma spraying (APS, LPPS, VPS, ...) to form a protective layer. Other methods can be considered to the same extent (PVD, CVD, SPPS, ...).

The protective layer 7 described also functions as an adhesive layer for superalloys.

On the protective layer 7 additional layers, in particular ceramic insulating layers 10 may be applied.

In the case of the component 1, the protective layer 7 is preferably applied on a substrate 4 made of nickel or cobalt based superalloy (FIG. 2).

Compositions of this type are known under the names GTD222, IN939, IN6203 and Udimet 500 as casting alloys. Further alternatives to the substrate 4 (FIG. 2) of the components 1, 120, 130, 155 are listed in FIG. 2.

The thickness of the protective layer 7 on the component 1 is preferably dimensioned to a value between about 100 μm and 300 μm.

The protective layer 7 is particularly suitable for protecting the parts 1, 120, 130, 155 against corrosion and oxidation, whereas the parts are in material temperature conditions of around 950 ° C. and about 1100 for aircraft turbines. Exposed to flue gas at material temperature conditions in and out of degrees Celsius.

Thus, the protective layer 7 according to the invention is a component of the gas turbine 100, in particular the guide vanes 120, the rotor that is exposed to the heating gas in front of or within the gas turbine 100 or the turbine of the steam turbine. It is particularly suitable for protecting the blade 130 or the heat shield member 155.

The protective layer 7 can be used as an overlay (protective layer is an outer layer) or as a bond coat (protective layer is an intermediate layer).

1 shows the layer system 1 as a component. The layer system 1 comprises a substrate 4. Substrate 4 may be metal and / or ceramic. In particular, turbine components such as turbine rotor blades 120 (FIG. 4) or turbine guide vanes 130 (FIGS. 3, 4), heat shield members 155 (FIG. 5), and steam or gas turbines ( In the case of further housing members (FIG. 3) of 100, the substrate 4 contains a nickel or cobalt or iron based superalloy, and in particular consists of the superalloy. Preferably, nickel based superalloys (FIG. 2) are used.

On the substrate 4, a protective layer 7 according to the invention is provided. Preferably, the protective layer 7 is applied by plasma spray (VPS, LPPS, APS, ...). The protective layer can be used as an outer layer (not shown) or as an intermediate layer (FIG. 1).

Preferably, a ceramic thermal insulation layer 10 is provided on the protective layer 7.

Preferably, the layer system consists of a substrate 4, a protective layer 7, a ceramic thermal insulation layer 10, and optionally a TGO under the thermal insulation layer 10.

The protective layer 7 can be applied on the newly manufactured parts and on the refurbished parts. Regeneration (modification) means that the parts 1 are optionally separated from the layers (insulation layer) after their use, and corrosion and oxidation products are removed, for example by acid treatment (acid stripping). In some cases, the cracks must also be repaired. After that, the part can be recoated again because the substrate 4 is very expensive.

In FIG. 3, the gas turbine 100 is shown in partial longitudinal section as an example. The gas turbine 100 includes a rotor 103 having a shaft 101 therein and rotatably mounted about a rotation shaft 102, which is also referred to as a turbine rotor. Along the rotor 103, for example, an annular combustion chamber 110, in particular an annular combustion chamber, and a turbine having an intake housing 104, a compressor 105, and a plurality of burners 107 arranged coaxially. 108 and the exhaust housing 109 are disposed. The annular combustion chamber 110 is connected with a heating gas channel 111 that is annular, for example. Here, for example, four turbine stages 112 connected in series form the turbine 108. Each turbine stage 112 is formed of, for example, two blade rings. Viewed in the flow direction of the working medium 113, within the heating gas channel 111, a row 125 formed of the rotor blades 120 is located, following the guide vane row 115.

In this case, the guide vanes 130 are fixed on the inner housing 138 of the stator 143, whereas the rotor blades 120 of the row 125 are for example rotor 103 by the turbine disk 133. ) Is mounted on. On the rotor 103 is a generator or operating machine (not shown).

During operation of the gas turbine 100, the air 135 is sucked and compressed by the compressor 105 through the intake housing 104. Compressed air supplied at the turbine side end of the compressor 105 is directed to burners 107 and mixed with fuel at the burners. The mixer is then combusted in combustion chamber 110, forming working medium 113. From the combustion chamber the working medium 113 flows along the heating gas channel 111 via the guide vanes 130 and the rotor blades 120. The working medium 113 expands while transmitting momentum in the rotor blades 120, whereby the rotor blades 120 drive the rotor 103 and the rotor drives an operating machine connected to the rotor itself.

Components exposed to the hot working medium 113 are exposed to thermal load during operation of the gas turbine 100. The guide vanes 130 and the rotor blades 120 of the first turbine stage 112, when viewed in the flow direction of the working medium 113, are the largest thermal next to the heat shield members lining the annular combustion chamber 110. Exposed to load. In order to withstand the temperature present beside the heat shield members, the temperature may be cooled by a coolant. To the same extent, the substrates of the components can also have a directional structure, that is to say that the substrates are single crystals (SX structures) or contain only longitudinally oriented particles (DS structures). For example, iron, nickel or cobalt based superalloys are used as materials for the components, in particular for the turbine blades 120, 130 and for the components of the combustion chamber 110. The superalloys are known, for example, from EP 1 204 776 B1, EP 1 306 454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949.

The guide vanes 130 include a guide vane root (not shown here) facing the inner housing 138 of the turbine 108 and a guide vane head positioned opposite the guide vane root. The guide vane head is fixed on the stationary ring 140 of the stator 143 while facing the rotor 103.

In FIG. 4, a rotor blade 120 or guide vane 130 of a turbomachine extending along the longitudinal axis 121 is shown in perspective view.

The turbomachine may be a gas turbine of an aircraft, or a gas turbine of a power plant for electric power generation, a steam turbine, or a compressor.

The blades 120, 130 comprise a fixed area 400 continuously along the longitudinal axis 121, a blade platform 403 adjacent to the fixed area, a blade surface 406, and a blade tip 415. .

As guide vanes 130, vanes 130 may include additional platforms (not shown) at vane tip 415.

The fixed region 400 is formed with a blade root 183 which is used (not shown) to secure the rotor blades 120, 130 on the shaft or disk. The blade root 183 is formed, for example, as a hammer head. Still other configurations as fir or swallowtail roots are possible. The blades 120, 130 include leading edges 409 and trailing edges 412 for the media flowing through the blade surface 406.

In the case of conventional blades 120, 130, for example, solid metal materials, in particular superalloys, are used in all areas 400, 403, 406 of the blades 120, 130. The superalloys are known, for example, from EP 1 204 776 B1, EP 1 306 454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949. In this case, the blades 120, 130 can be produced by a casting method, also by directional solidification, by a forging method, by a milling method, or by a combination thereof.

Workpieces containing single crystal structures or structures are used as components for machines that are exposed to high mechanical, thermal and / or chemical loads during operation. The production of single crystal workpieces of this type is carried out, for example, by directional solidification of the melt. This is a casting method in which a liquid metal alloy is solidified in a single crystal structure, that is, a single crystal workpiece, or directionally. In this case, the dendritic crystals are oriented along the heat flow, forming columnar crystal grain structures (particles extending over the entire length of the workpiece, ie, referred to herein as directional solidified according to general terms), It forms a single crystal structure, that is, the entire work piece consists of a single crystal. In these methods, the transition to isotropic (polycrystalline) coagulation should be avoided, because non-directional growth necessarily results in the formation of transverse and longitudinal grain boundaries, which result in the directional solidification or of the monocrystalline part. This is because the superior characteristics are invalidated. Thus, generally speaking a directional solidified microstructure, it possesses longitudinally extending grain boundaries, as well as single crystals that do not have grain boundaries or at best have small grain boundaries, but transverse grain boundaries It also means columnar crystal structures that do not possess. The latter crystal structures are often described as directionally solidified structures. Methods of this type are known from US-PS 6,024,792 and EP 0 892 090 A1.

To the same extent, the blades 120, 130 may comprise protective layers 7 according to the invention against corrosion or oxidation. The density is preferably 95% of the theoretical density. A protective aluminum oxide layer (TGO = thermal growth oxide layer) is formed on the MCrAlX layer (either as an intermediate layer or as the outermost layer).

On the MCrAlX still preferably may be provided with an outermost layer, a heat insulating layer, the insulating layer is, for example consists of _{ZrO 2, Y 2 O 3 -ZrO} 2, That is, the heat insulating layer is yttrium oxide and / or calcium oxide and / or magnesium oxide Is not stabilized, partially stabilized, or completely stabilized. The insulating layer covers the entire MCrAlX layer. Columnar grains are created in the thermal insulation layer through suitable coating methods such as, for example, electron beam physical vapor deposition (EB-PVD). Other coating methods can also be considered, such as atmospheric plasma spray (APS), LPPS, VPS or CVD. The thermal insulation layer may include particles that are porous for further improved thermal shock resistance or that retain micro cracks or macro cracks. In other words, the thermal insulation layer is preferably more porous than the MCrAlX layer.

The blades 120 and 130 may be formed in a hollow or solid shape. If the blades 120, 130 are to be cooled, the blade is hollow and optionally also includes film cooling holes 418 (shown in broken lines).

In FIG. 5, the combustion chamber 110 of the gas turbine 100 is shown. The combustion chamber 110 is formed as a so-called annular combustion chamber, for example, in which a plurality of burners 107 arranged in the circumferential direction about the rotational shaft 102 generate a flame 156 while passing through a common combustion chamber space 154. . To this end, the combustion chamber 110 is formed as an annular structure which is positioned about the axis of rotation 102 as a whole.

In order to achieve a relatively high efficiency, the combustion chamber 110 is configured for a relatively high temperature of the working medium M at about 1000 ° C to 1600 ° C. In order to allow for a relatively long operating period even under conditions of undesirable operating parameters for the materials, the combustion chamber wall 153 is provided with heat shield members 155 on its face towards the working medium M. FIG. It has an inner lining formed into.

Further, based on the high temperature inside the combustion chamber 110, a cooling system may be provided for the heat shield members 155, or for holding members of the heat shield members. In this case, the heat shield members 155 are also hollow and optionally include cooling holes (not shown) which pass through into the combustion chamber space 154.

Each heat shield member 155 made of an alloy has a particularly heat resistant protective layer (MCrAlX layer and / or ceramic coating layer) on the working medium side or is made of a high temperature resistant material (solid ceramic stone). The protective layers 7 may be similar to the case of turbine blades. Still may be for example provided with a ceramic heat insulating layer formed on the MCrAlX, a heat insulating layer is for example, ZrO _2, Y ₂ O ₃ -ZrO ₂ is composed of, i.e. a heat insulating layer is unsubstituted or stabilized by yttrium oxide and / or calcium oxide and / or magnesium oxide , Partially or completely stabilized. Columnar grains are produced in the thermal insulation layer through suitable coating methods such as, for example, electron beam physical vapor deposition (EB-PVD). Other coating methods can also be considered, such as atmospheric plasma spray (APS), LPPS, VPS or CVD. The thermal insulation layer may include particles that are porous for further improved thermal shock resistance or that retain micro cracks or macro cracks.

Regeneration (modification) means that after the use of the turbine blades 120, 130 or the heat shield members 155, the protective layers have to be removed from these blades or heat shield members as the case may be (eg through sand blasting). do. Thereafter, removal of the corrosion layer and / or the oxidation layer and the product thereof is carried out. In some cases, cracks in the turbine blades 120 and 130 or the heat shield member 155 are also repaired. Thereafter, the turbine blades 120, 130, or heat shield members 155 may be recoated and the turbine blades 120, 130, or heat shield members 155 may be used again.

Claims (12)

Alloy,
At least elements (in weight%), ie
22% to 24% cobalt (Co), in particular 23% cobalt (Co),
14% to 16% chromium (Cr), in particular 15% to 16% chromium (Cr), very particularly 15% chromium (Cr),
10.5% to 11.5% aluminum (Al), in particular 11% by weight of aluminum (Al),
At least one metal, in particular yttrium (Y), selected from the group comprising elements of scandium (Sc) and / or rare earths of 0.2% to 0.4%, in particular 0.3%,
Optionally 0.3% to 0.9% tantalum (Ta), in particular 0.4% to 0.7% tantalum (Ta), very particularly 0.5% tantalum (Ta),
An alloy containing nickel (Ni), in particular nickel (Ni) balance. The alloy of claim 1 containing 0.3 wt% yttrium (Y). The alloy of Claim 1 or 2 which does not contain rhenium (Re). The alloy of claim 1, which does not contain silicon (Si). The method according to any one of claims 1 to 5,
Containing tantalum (Ta),
Especially containing 0.4% by weight of tantalum (Ta),
Very particularly 0.5% by weight of tantalum (Ta). 6. The method according to any one of claims 1 to 5,
Free of zircon (Zr) and / or free of titanium (Ti) and / or free of gallium (Ga), and / or free of germanium (Ge), and / or platinum ( Free of Pt) and / or free of hafnium (Hf) and / or free of cerium (Ce) and / or free of iron (Fe) and / or palladium (Pd) An alloy which does not contain and / or does not contain boron (B) and / or does not contain carbon (C). The alloy according to claim 1, which is composed of cobalt, chromium, aluminum, yttrium, nickel and tantalum as an optional component. The alloy according to claim 1, which is composed of cobalt, chromium, aluminum, yttrium, nickel, and tantalum. The alloy of claim 1, wherein nickel (Ni) forms a matrix. Especially at high temperatures,
A protective layer 7 for protecting the component 1 against corrosion and / or oxidation,
Protective layer (7) having a composition of the alloy according to at least one of the preceding claims. 11. The method of claim 10,
Plasma sprays, especially APS or
Protective layer 7 applied by high speed spray (HVOF). part,
In particular the components 120, 130, 155 of the gas turbine 100,
In particular, the substrate 4 of the components 120, 130, 155 is based on nickel, cobalt based,
The substrate comprises a protective layer 7 according to claim 10 or 11 for protection against corrosion and oxidation at high temperatures,
In particular, the ceramic heat insulation layer 10 is applied on the protective layer 7.

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